Aircraft with integrated lift and propulsion system

ABSTRACT

A vertical take-off and landing (VTOL) aircraft is designed to be so efficient that it can be commercially competitive with runway dependent aircraft operating in a range of 100 to 1000 miles or more, with glide ratios of at least 24. Improvements include a combination of high efficiency tilting rotor and wing design that enable both vertical takeoff and efficient high speed cruising, a high aspect ratio wing, and a variable speed propulsion system efficient in both hover and cruise flight. Preferred aircraft use narrow chord inboard and outboard wings, and use efficient lightweight design to achieve unusually low empty weight fraction. In some embodiments, the rotors use medium rather than high modulus fibers, with wider blades of lower taper ratio, to provide the stiffness and mass properties required for high performance OSTR rotor blades. Also disclosed are VTOL aircraft with glide ratios from approximately 26 to over 40.

This application is a continuation-in-part (CIP) of U.S. applicationSer. No. 14/010382 filed Aug. 26, 2013, which is a divisional of U.S.application Ser. No. 12/857,191 filed Aug. 16, 2010, which is adivisional of U.S. application Ser. No. 12/429,990 filed Apr. 24, 2009which is Continuation-In-Part of U.S. application Ser. No. 11/505,067filed Aug. 15, 2006, which claims priority to U.S. ProvisionalApplication No. 60/708,805 filed Aug. 15, 2005, and further claimspriority to U.S. Provisional Application No. 61/047,844 filed Apr. 25,2008, each of which is incorporated by reference herein in its entirety.

FIELD OF THE INVENTION

The field of the invention is high efficiency vertical takeoff aircraft.

BACKGROUND

Passenger transit can be broken down into categories based on the lengthof the trip as shown in FIG. 1. For short trip distances, under 2 miles,the transit modes of walking 110, and bicycling 120 are preferred: theyrequire very little overhead (non-travel time associated with the trip,such as putting on shoes) but are very slow. For somewhat longer trips,between 2 and 200 miles, driving 130 is the preferred mode of travel.However, external factors such as rising fuel prices and increasingtraffic congestion are driving down the speed of this travel mode andincreasing its cost, especially relevant for trips over 100 miles.Fixed-wing air travel 140 offers high trip speeds, and is usuallypreferred for very long distances, over 1000 miles. However, there is alarge overhead associated with this mode of travel, including travel toan airport, airport security lines, waiting for takeoff, and groundtransport at the destination, all of which reduces the efficiency oftraditional fixed-wing air transport for mid-range, or regional, tripsbetween 100 and 1000 miles. As a result, there is a largely unmet needfor fast, affordable regional transport 150.

Although the prior art seems to have appreciated the need for fast,affordable regional transport, the focus has mostly been on railtransit. Rail travel typically offers higher average speeds than cartravel, but is constrained to operate within the bounds of fixed railinfrastructure. For regions with a dispersed population, includingsuburbanized regions, the cost of building rail infrastructure toconnect a large percentage of the population is prohibitive.

VTOL (vertical take-off and landing) transport has also been attemptedover the years for regional transport, and has clear potentialadvantages over rail travel. The ability to take off or land verticallyenables passengers to start and finish their journey near their trueorigin or destination, be it an urban center or intersection offreeways. However, despite decades of attempts, successful VTOLtransport remains elusive due to challenging technical obstacles. Amongother things, prior art VTOL is prohibitively expensive to operate, haslow flight speeds, limited ranges, is relatively fuel inefficient, andhas a relatively poor safety record. This history is catalogued to someextent in the books “The Principles of Helicopters Aerodynamics”, J. G.Leishman, 2006 and “The Helicopter: Thinking Forward, Looking Back”, J.G. Leishman, 2007. Technical details about individual aircraft can beobtained from the book series “Jane's All the World's Aircraft” byreferencing the appropriate volume.

These references, as well as all other extrinsic materials discussedherein, are incorporated by reference in their entirety. Where adefinition or use of a term in an incorporated reference is inconsistentor contrary to the definition of that term provided herein, thedefinition of that term provided herein applies and the definition ofthat term in the reference does not apply.

In the void left by the lack of a viable regional transportation method,passengers are relegated to driving long distances in increasinglycongested traffic, or flying regional fixed-wing transport aircraft andenduring long airport-related waits. There is little or no appreciationin the prior art that vertical take-off aircraft should or even could bemodified to simultaneously have high flight speed, high fuel efficiency,and the ability to carry a substantial payload.

FIG. 2 shows a typical prior art tiltrotor aircraft 200 comprising awing 202, a fuselage 204, and a first tilting rotor system 210comprising a first rotor blade 212 and first nacelle 218 in aircraftcruise mode corresponding to a generally horizontal position of thenacelle 218. The aircraft is also equipped with a second tilting rotorsystem 220 on the opposite end of the wing 202. The second rotor system220 is depicted in conversion from a horizontal position consistent withaircraft cruise mode to a vertical position consistent with helicoptermode.

In practice, nacelles 218, 228 on either side of the aircraft in priorart tiltrotors have a substantially identical tilt angle. The tilt angle236 of a nacelle 228 is the angle 236 between the tilting nacelle axis238 and the aircraft axis 234. In a typical tilt rotor aircraft 200, thenacelle 204 is also capable of operation in a generally verticalposition used in helicopter mode flight. The nacelle 228 tilt angle 236is usually affected using a tilt actuator and mechanism to convert fromhelicopter mode flight to aircraft cruise mode. A cross-shaft 206 isdisposed within the wing 202 and runs between left and right nacelles218, 228.

The aircraft of FIG. 2 has a relatively small fuselage, a relativelysmall wing aspect ratio, and has gimbaled rotors not capable of variablespeed operation. It appears to have been optimized for thethen-contemplated use of short-range transport with relatively fewpassengers, under which fuel efficiency and speed are not so importantas vertical take-off capability and cost. There is nothing in the priorart indicating that those of ordinary skill in the art appreciated thatvertical take-off aircraft should, or even could, be optimized for highfuel efficiency, while carrying a substantial payload.

The majority of prior art rotorcraft and prior art vertical takeoffaircraft are conventional helicopters. Conventional helicopters, such asthe modern Sikorsky™ S-92, are severely limited in terms of cruise speedand efficiency. A conventional helicopter is lifted and propelled by thesame predominantly horizontal rotor or rotor, one side of which advancesinto the oncoming flow, and one side of which retreats away from it.During cruise, the airspeed towards the tip the advancing rotor blade ismuch higher than that of the helicopter itself. It is possible for theflow near the tip blade to achieve or exceed the speed of sound, andthus produce vastly increased drag and vibration. This limits theforward speed of helicopters. Additionally, a rotor is an inefficientway to generate lift as compared to a wing, partially due to thedissymmetry of lift between advancing and retreating sides of the rotor.

A major step forward in the prior art was the tiltrotor configurationincluding, for example, the Bell™/Augusta™ BA609. Tilt-rotors representa major step forward because they generate most or all of the liftnecessary for cruise flight with a wing instead of rotors, which isconsiderably more efficient than rotor borne flight. Prior arttilt-rotors have had short, low-aspect ratio wings that were relativelythick because they had to support heavy rotor systems, which results inlower efficiency, L/D, as compared to fixed-wing aircraft.

Despite marginal increases in speed and forward flight efficiency oftilt-rotor aircraft relative to helicopters, the prior art tilt-rotorshave failed to improve on the productivity (how fast one transports apayload) of conventional helicopters. This is because the ability of amodern tiltrotor to cruise up to 50% faster than a modern helicopter isalmost entirely offset by its relatively higher empty weight fraction(aircraft empty weight divided by maximum hover takeoff weight,typically around 0.60-0.65 for prior art tilt-rotor) as compared toconventional helicopters.

Unless a contrary intent is apparent from the context, all rangesrecited herein are inclusive of their endpoints, and open-ended rangesshould be interpreted to include only commercially practical values.Similarly, all lists of values should be considered as inclusive ofintermediate values unless the context indicates the contrary.

The complexity, resulting high cost, aerodynamic inefficiency, poorsafety record, and weight criticality of rotorcraft conspire to makethem entirely uncompetitive with fixed-wing aircraft. Modern prior artrotorcraft have a productivity, the product of payload carried andspeed, about 5 times lower than modern turboprop airliners, as reportedin the AHS Journal paper “Rotorcraft cost too much” by Harris andScully.

In view of the existing prior art and present state of technology, andwithout access to the present inventive subject matter, a person ofordinary skill in the art attempting to design an advanced futuretiltrotor transport would tend to make a smaller diameter conventionalrotor to minimize the rotor system weight. This would result in higherrotor disc loading in hover, and would constrain the wing to be shortand thick to prevent the whirl flutter aeroelastic instability resultingfrom heavy conventional rotors at higher speeds. Furthermore, the flightenvelope would be constrained to lower speeds because of the thick wing,the efficiency would be reduced because of the short wing, and thecruise altitude would be capped by the small wing area.

In summary, what are still needed are: (1) an appreciation that VTOLtransport systems could realistically be as efficient and productive asfixed wing aircraft for regional transport of substantial payloads; and(2) technologies that could be used to implement such systems. In orderto implement those goals, an aircraft would realistically need to havesome or all of the following characteristics:

-   -   a. Tailored, efficient aerodynamics, especially including the        inner wing, rotor blade and nacelle shaping, for efficient        cruise flight at speeds up to Mach 0.65 (100 knots faster than        the prior art Bell™ V-22);    -   b. Wing area and wing airfoil technology (expressed as an        M²C_(L) of at least 0.30-0.35) to provide for cruise flight at        35,000-41,000 feet (above most adverse weather and 10,000-16,000        feet higher than prior art tilt rotor aircraft);    -   c. Small empennage, low drag fuselage, low drag landing gear        fairing and high aspect ratio wing to assist in providing a        lift-to-drag ratio of 16-23 (3-4 times higher than the prior art        Bell™ V-22);    -   d. Structures that support a very low aircraft empty weight        fraction while sustaining rotor loads and providing the        strength, stiffness, and durability of a high speed pressurized        commercial transport (empty weight fractions 20-40% lower than        in the non-pressurized prior art Bell™ V-22);    -   e. Mechanical systems that support the low aircraft empty weight        fraction while sustaining rotor loads, and providing needed        functionality;    -   f. A rotor system that provides a high cruise flight propeller        efficiency at a cruise Mach number of 0.65 while also being        capable of vertical takeoff; and    -   g. Low hover download to reduce the amount of engine power        required for hover. (less than 5% of rotor lift vs. 11% the        prior art Bell™ V-22 tilt-rotor).

The parent application (Ser. No. 14/010382) described systems that wouldcarry at least 20,000 pounds of payload, have a wing sized anddimensioned to have a maximum wing loading of between 60 and 140 poundsper square foot, have a wing aspect ratio between 10 and 22, be capableof cruise lift-to-drag ratio of between 13 and 26, and be capable ofsustained cruise flight with the first rotor operating at a rotationalspeed no greater than 75% of the operational maximum rotational speed.

The rotor blades of such systems could be manufactured using highmodulus graphite fibers, utilizing techniques described in Aircraft WithIntegrated Lift And Propulsion System (U.S. Pat. No. 7,861,967), andpossibly utilizing techniques described in Wing and Blade StructureUsing Pultruded Composites (U.S. Pat. No. 8,114,329), including the useof high modulus graphite pultrusions in the critical regions of theblade structure. The latter method offered an increase in compressionstrength over composite plies of the same fiber, while still maintainingthe high stiffness to weight ratio needed for OSR and OSTR rotordynamics.

That high modulus graphite provides high blade stiffness and low bladeweight needed to avoid aeroelastic instability was taught in OptimumSpeed Tilt-Rotor (OSTR) (U.S. Pat. No. 6,641,365) and Optimum SpeedRotor (OSR) (U.S. Pat. No. 6,007,298).

What is not apparent from the prior art or the parent application,however, is how to achieve the desired effects of OSR and OSTR usingmedium modulus graphite fibers instead of a high modulus graphitefibers. What is also not apparent from the prior art or the parentapplication, is how to construct a rotorcraft capable of hover takeoff,and having a glide ratio of greater than 40 at 0.55 Mach.

SUMMARY OF THE INVENTION

The advantages of medium modulus fibers over high modulus fibers arewell known, the medium modulus carbon fibers offer substantially higherstrength to failure, 2-3 fold higher elongation to failure and as aresult are less fragile. The universal acceptance of medium modulesfibers in advanced aircraft is also motivated by their substantiallylower cost. What is taught in this application is how to make variableRPM rotor blades with medium modulus fibers, as opposed to high modulusfibers, and still achieve the critical rotor aero-structural dynamicstaught with respect to OSR and OSTR blades.

In one aspect, the inventive subject matter provides apparatus, systemsand methods in which a rotorcraft is flight operational at variable RPM,where the spar(s) of the rotor blades have a high percentage of mediummodulus fibers.

Preferred such rotorcraft have relatively large diameter rotors, up to36 feet or more, with one or more relatively deep spars, and maximumroot chord of at least 8% of rotor diameter. Some contemplatedembodiments allow the rotor to operate at a variable RPM correspondingto a rotational tip of 0≦Mach number≦0.75 Mach, while also lift loadedbetween 0 lift coefficient and stall. The average elongation to failureof the medium modulus fibers is preferably at least 1.5%, morepreferably at least 1.8%, and most preferably at least 2%.

In another aspect, the inventive subject matter provides apparatus,systems and methods in which a tilt rotor aircraft capable of hovertakeoff has a cross sectional area 3-15% of a planform area of the fixedwing, and still have a glide ratio of 26 to over 40 at 0.55 Mach.

Preferred such rotorcraft is flight operational at variable RPM, wherethe spar(s) of the rotor blades have a high percentage of medium modulusfibers. Preferred such aircraft also have a large ratio of wing span torotor diameter, of at least 2.5 or even at least 2.6. By way ofcomparison, the TR53 taught in the parent U.S. Pat. No. 8,517,302 has aratio of wing span to rotor diameter of 2.415). Rotor diameters ofgreater than 30 feet are contemplated, including rotors of at least 32feet, and up to 36 feet or more.

In some embodiments the aircraft's fuselage has a relatively smallmaximum cross-sectional area in comparison with the planform area of thefixed wing. But fuselage maximum cross-sectional area can vary amongdifferent embodiments, with correspondingly different glide ratios.

All suitable power plants, gearboxes, and drive trains are contemplated.Especially preferred aircraft have a turbine engine sized anddimensioned to provide sufficient power to hover the aircraft, and alsoovercome an aerodynamic drag of the aircraft even in high-speedhigh-altitude cruise flight mode. As a demonstration, some embodimentsof high glide ratio rotorcraft can circumnavigate the globe withoutrefueling and break the unrefueled range world record and the speed incircumnavigation record currently held by a jet-powered aircraft(GlobalFlyer) after taking-off in hover out of ground effect.

Various objects, features, aspects and advantages of the inventivesubject matter will become more apparent from the following additionaldescription of preferred embodiments, along with the accompanyingdrawings in which like numerals represent like components.

BRIEF DESCRIPTION OF THE DRAWING

FIG. 1 is a graph depicting relative uses of different modes of travel.

FIG. 2 is a perspective view of a typical prior art tiltrotor aircraft.

FIG. 3 is a side view of a preferred embodiment in flight, showingforces acting on the aircraft.

FIG. 3A is a schematic cross-section of a wing of the aircraft of FIG.3.

FIG. 3B is a schematic cross-section of a rotor of the aircraft of FIG.3.

FIG. 4 is a perspective view of another preferred aircraft.

FIGS. 5, 6 and 7 are top, side and front views, to scale, of anespecially preferred aircraft.

FIG. 8 is a table of selected dimensions of the aircraft of FIGS. 5-7.

FIG. 9 is a perspective view of a preferred nacelle structure.

FIG. 10 is exploded view of a preferred aircraft structure.

FIG. 11 is a table containing preferred weights as calculated for twoversions of the TR53 as verified using the analyses described above.

FIG. 12 is a schematic view of a preferred rotor blade showing airfoilsection thicknesses at various stations on the blade.

FIG. 13 is a schematic of a preferred nacelle.

FIGS. 14A and 14B are vertical cross sections of alternative wingairfoils.

FIG. 15 is a graph showing power reduction from reducing download orincreasing wing aspect ratio.

FIG. 16 is a schematic of a preferred aircraft landing gear.

FIG. 17 is a graph showing efficiency of a preferred rotor and showingthat high efficiency can be maintained by varying rotor speed.

FIGS. 18, 19A-19F and 20 are top, side and front views, to scale, of anespecially preferred rotor blade of the current invention

FIGS. 21 and 22 are tables depicting design data for the blade of FIGS.18, 19A-19F and 20.

FIG. 23 is a cross-section of an airfoil of a preferred embodiment.

FIG. 24 is a table depicting sample carbon fiber flight thickness andorientation in cell C along the blade of FIGS. 18, 19A-19F and 20.

FIG. 25 is a graph depicting mass/length as a function of radial stationof the blade of FIGS. 18, 19A-19F and 20.

FIG. 26 is a graph depicting stiffness as a function of radial stationof the blade of FIGS. 18, 19A-19F and 20.

FIG. 27 is a graph depicting deflection response as a function of radialstation of the blade of FIGS. 18, 19A-19F and 20.

FIG. 28 is a graph depicting natural frequency as a function rotor RPM(fan plot) of the blade of FIGS. 18, 19A-19F and 20.

FIG. 29 is a graph depicting mode damping ratio in hover as a functionof Rotor CT/σfor a preferred aircraft using a rotor with blades of FIGS.18, 19A-19F and 20.

FIG. 30 is a graph depicting damping ratio as a function of true airspeed for a preferred aircraft using a rotor with blades of FIGS. 18,19A-19F and 20.

FIG. 31A is a top and front view to scale, of a preferred highlift-to-drag ratio VTOL aircraft configuration. FIG. 31B shows aperspective view of the same configuration. FIGS. 31C and 31D show bothside views, with the tilt rotor in different orientations.

FIG. 32 is a table depicting design data for the aircraft of FIG. 31.

FIG. 33 is a graph and associated data table showing Lift/Drag (Glide)ratio as a function of fuselage cross-sectional area and fuselage lengthas a function of wing area of the aircraft configuration of FIG. 31.

DETAILED DESCRIPTION

Physics of Efficient Flight

As shown in the side view illustration of FIG. 3, a preferred aircraft300 in non-accelerated flight is in force equilibrium. Its weight 302 isbalanced by the same amount of lift 304, and the aerodynamic resistanceto its motion, drag 306, is countered by a propelling thrust 308. Anaircraft in steady level equilibrium flight is said to be in cruiseflight.

The generation of thrust requires energy, often produced via combustionin a turbine engine. Efficient flight minimizes the amount of energyadded, or equivalently, the amount of fuel burned to maintainequilibrium flight.

FIG. 3A provides detail of the wing 310 while FIG. 3B provides detail ofthe rotor 340.

A flying machine must generate lift 304 and thrust 308 forces to fly.Lift is generated by air pressure differences across lifting surfaces,especially wings 310. However, lifting surfaces also generate some drag,including induced drag, due to the direction of the generatedaerodynamic force. They, like other parts of an aircraft, also generatedrag due to friction and viscous effects. Wings that are long andnarrow, said to be of high aspect ratio, are more efficient than shortand wide wings in that they produce more lift for less drag (called thelift-to-drag ratio or L/D). However, there are structural limits on howlong and narrow a wing can be, especially if the wing structurallysupports an engine or rotor. In the case of a helicopter, or a tiltrotorin helicopter mode, lift is generated by the rotor instead of by wings.

It is further advantageous to minimize drag, and thereby the thrust thatis generated for flight, by judiciously designing and streamlining thenon-wing portions of the aircraft, including the fuselage 330. Anefficient aircraft configuration, including for example a sailplane,might have a small fuselage diameter to minimize drag and a long highaspect ratio wing. The overall aerodynamic efficiency of an aircraft canbe expressed as an aircraft lift-to-drag ratio.

An aircraft also produces thrust 308 sufficient to counter drag 306.This is often accomplished by the use of a turbine engine attached to afan in the case of a turbofan or jet engine, a propeller in the case ofa turboprop, an inclined rotor in the case of a helicopter, or aprop-rotor 340 in the case of a tilt-rotor. For the most efficientgeneration of thrust, a large quantity of air is accelerated by therotor 340 to a speed only slightly higher than the flight speed. Thismaximizes the propulsive thrust generated while minimizing the energyadded. This principle is responsible for the relatively higherefficiency of turboprop engines as compared to turbofan engines atmoderate flight speeds. The propulsive efficiency of a propeller,prop-rotor, or rotor is expressed as η, the cruise propeller efficiency,the product of thrust generated by the propeller and aircraft forwardspeed, divided by the power input to turn the propeller.

Helicopters generate both lift and thrust forces using rotors. In thecase of a hovering helicopter, the thrust and lift forces are equivalentbecause there is no vehicle forward motion. Thus, for helicopters, theefficiency of generating thrust is particularly important. Forefficiency, larger rotors are preferred. However, larger rotors areheavier, and their structural supports tend to constrain rotor diameter.Weight is of particular concern for helicopters, which are often said tobe weight critical, because the helicopter rotor alone must generatesufficient thrust to become airborne. In forward flight, a rotor has alow L/D as compared to a fixed wing generating the same lift. Thus, inforward flight, conventional helicopters are inefficient as compared towing-borne or fixed-wing aircraft. The efficiency of a hovering rotor isquantified by the figure of merit, or FM, which is the ratio of idealinduced power to actual power required to turn the rotor. The bestpossible FM is 1.0, a typical rotor FM is between 0.60 and 0.80. Thedisc loading, DL, of rotor is defined as the ratio of thrust generatedby a rotor in pounds, T, divided by rotor area in square feet.

To compare the relative efficiency of jet-powered aircraft,propeller-driven aircraft, and helicopters, it is useful to introducethe parameter L/De, the effective lift-to-drag ratio, the lift anaircraft produces divided by its propelling force. For a jet aircraft,L/De is simply equal to L/D. For a propeller-driven aircraft ortilt-rotor, L/De is equal to the product of aircraft configurationefficiency, L/D, and the cruise propeller efficiency, η. To illustratethe startling gap in efficiency between prior art rotorcraft andfixed-wing transports, consider that a modern efficient transporthelicopter typically has an L/De between 4 and 5, while a modern jettransport has an L/De between 14 and 20.

FIG. 4 is a perspective illustration of a preferred tiltrotor aircraft400 according to the present teachings, comprising a wing 402, cockpit406, and fuselage 404, a first tilting rotor system 410 shown inhelicopter-mode position, and a second tiling rotor system 450 shown inairplane cruise-mode position. These rotor systems have 75-foot diameterrotors. In practice, both the first and second tilting rotor systems410, 450 are likely to have substantially the same orientation in flightat any given time. A rotor system 450 comprises rotor blades 456, 458that trace a path 464 that defines a rotor diameter 462. Preferredrotors have rotor diameters of at least 20 feet, 40 feet, 53 feet, 65feet, 75 feet, 90 feet, or even 120 feet.

In this figure, the wing 402 remains geometrically fixed to the fuselage404 during flight in both vertical takeoff mode or cruise flight.However, folding, rotating, and tilting wings are also contemplated. Inpreferred embodiments, the first and second rotors in the first andsecond rotor systems 410, 450 are of a stiff hingeless variety,including for example that described in U.S. Pat. No. 6,641,365. Such arotor system 450 transmits considerable forces and moments to the wing402 and fuselage 404. Hingeless rotor systems for tiltrotors are unlikeprior art gimbaled systems in that they can transmit considerable largemoments, also referred to as mast moments, to the airframe. Inairplane-mode cruise flight, the rotor system 450 generates thrust asindicated by block arrow 452 and moment as indicated by block arrow 454.A rotor system 410 in helicopter mode also generates thrust and moments(not shown).

Rotor system 410 comprises a tilting nacelle 418, which also serves as atilting mast in the case of this hingeless rotor system, and a hub 440that is not gimbaled with respect to the nacelle 418. The rotor rotatesabout the hub axis 422, also known as the “rotational axis of the hub”.It can be seen that the rotor comprising rotor blades 414, 416 isdisposed on a mast such as the nacelle 418.

The tilt angle, indicated by arrow 432, is the angle between thehorizontal airframe axis 420 and the hub axis 422. The entire rotorsystem 410, including the nacelle 418 and hub 440, is tiltably coupledto the wing 402 by means of a tilt actuator and spindle. The rotorsystem 410 tilts with respect to the wing 402 about the tilt axis 424.Although the nacelle 418 or mast tilts, it is considered “non-rotating”structure. The term “rotating structure” refers to the hub 440, blades414, 416, spinner 412 and other components that rotate with the rotor.

First and second blades 414, 416 are preferably coupled to the hub 440without hinges in the flap direction 438 and lag direction 439. Theblades 414, 416 also transmit blade bending moments to the hub in theflap direction 438 and lag direction 439. In turn, the hub 440 transmitsthrust and large hub bending moments to non-rotating structure includingthe nacelle 418, wing 402, and fuselage 404. In preferred embodiments,blade bending moments of 40000, 70000, 100000, 300000 and even 500000foot-pounds are contemplated. Likewise, hub bending moments, asindicated by arrow 454, of at least 50000, 100000, 200000, 500000, andeven 1000000 foot-pounds are contemplated. The structure and design of apreferred hub 440 have elements that allow for the accommodation of suchvery large blade and hub moments.

The aircraft 400 is also equipped with a sponson 468, which serves as aprovision for storing fuel. Preferred aircraft have a fuel capacitysufficient to provide a maximum range of at least 500, 1000, 2000, oreven 5000 miles carrying its maximum payload. Some preferred aircrafthave provisions (not shown) for storing fuel in fuel tanks in the wing.

Layout and Dimensions of an Especially Preferred Aircraft

Among the contemplated aircraft is a most preferred embodiment (dubbedthe TR53) of a commercial passenger transport that is expected to beable to compete cost-effectively with jet transports and with high speedtrains. The TR53 is named as such because it has two 53 foot diameterside-by-side rotors. Those skilled in the art will appreciate that theinventive principles discussed herein are also applicable to a widerange of other rotorcraft, including those having larger or smallerrotors, and other rotor configurations.

FIG. 5 shows a top-view drawing of a TR53 aircraft 500, while FIG. 6 isa side view and FIG. 7 is a front view of the same aircraft. Theaircraft comprises a fuselage 504 sized and dimensioned to carry apayload. Preferred payloads include passengers, package freight, andother cargo. The TR53 is capable of carrying payloads of up to between26000 pounds and 42000 pounds depending on the configuration. Naturally,the aircraft can carry payloads less than its maximum payload, and thusthe fuselage is sized to carry at least 20000 pounds of payload.Contemplated fuselages for other aircraft designed according to theteachings herein can be sized to carry at least 5000, 10000, 50000, or82000 pounds.

The aircraft 500 is also advantageously equipped with first and secondrotors 510, 520, which are sized and positioned to be able to eitherlift the aircraft vertically or provide thrust in cruise flight. A firstrotor 510 is carried by a tilting nacelle 518, thereby making the firstrotor 510 a tilting rotor. The tilting nacelle 518, in turn, is carriedby an inboard wing section 502. The wing is sized and dimensioned tosupport the aircraft in cruise flight, and is disposed to carry thefirst and second tilting rotors 510, 520. In the TR53, the wingcomprises an inboard wing section 502, and first and second outboardwing sections 508, 509. The total wingspan 602 is the length from onewingtip to the other. The wing halfspan 622 is the distance from thefuselage centerline to a wingtip. The inboard wing section 502 has astation and a wing airfoil at mid span 624, which is half-way betweenthe fuselage centerline and a wing tip. The chord 626 of the inboardwing section 502 is the width of a section. At mid span 624, the wingsection has a chord and thickness. The wing is advantageously configuredto have a thickness-to-chord ratio of at between 19% and 22% or evenbetween 18% and 21% in order to provide for high-speed flight.

The aircraft 500, like any aircraft, has an empty weight when neitherfuel nor payload are present on an otherwise operational aircraft.Similarly, the aircraft 500 has a maximum takeoff weight, beyond whichthe aircraft no longer has the capability to become airborne under itsown power. The maximum takeoff weight is usually associated with amaximum combination of fuel and payload. For a verticaltakeoff-aircraft, an aircraft has a maximum hover weight, which is themaximum weight at which an aircraft can hover under its own power andmaintain structural integrity. The aircraft 500 can also perform ashort-takeoff, with a higher takeoff weight. While these weights canvary somewhat with atmospheric conditions, sea level altitude, standardatmosphere conditions are assumed herein unless otherwise specified.

The wing, comprising an inboard wing section 502 along with first andsecond outboard sections 508, 509, has a planform area, which is theprojected area of the wing when viewed from the top, as in FIG. 5. Inthe case of tilting wings, the planform area is the greatest projectedarea as the wing tilts. The wing also has a wetted area, which is thetotal exterior surface area of the wing, including the underside and topside of the wing. An aircraft having a wing has a maximum wing loadingthat is defined as the maximum takeoff weight divided by the wingplanform area. Preferred embodiments have maximum wing loadings between60 and 140 pounds per square foot.

A wing also has an aspect ratio defined as the square of the totalwingspan 602 divided by the wing planform area. Preferred aircraft havewings that are proportioned to have an aspect ratio between 10 and 22,12 and 20, or 14 and 18. Achieving these high aspect ratios is difficultin tiltrotors because the wing must support tilting rotors 510, 520.Further, a wing has a leading edge sweep angle 628, which is defined tobe the average angle between a line normal to the fuselage centerlineand the leading edge of the inboard wing section 502 that goes betweenthe fuselage 504 and a rotor 520. Preferred aircraft have leading edgesweep angles of less than 20°, 15°, 10°, or even 5°.

As shown in FIG. 5 and FIG. 6, the TR53 also has an empennage 570,comprising a vertical tail 572 and a horizontal tail 574. The horizontaltail 574 has a span 674 and the vertical tail 572 has a height 672. Theempennage 570 has a wetted area, defined as the total exterior surfacearea of all tail surfaces, including the vertical tail 572 andhorizontal tail 574. The TR53 empennage 570 has a wetted area of 372square feet, while the wing has a wetted area of 2092 square feet and aplanform area of 1000 square feet. Preferred aircraft have an empennage,where the wetted area of the empennage is between 14 and 40 percent, oreven between 10 and 20 percent of the wetted area of the wing.

The TR53 aircraft 500 has a fuselage 504, the payload-carrying portionof which has a width 604 and a height 606 as shown in FIGS. 5 and 6. Asshown in FIG. 7, the fuselage 504 has a frontal area, defined as theprojected area of the fuselage when viewed from the front. The TR53fuselage has a frontal area of 133 square feet. Preferred aircraft havea frontal area that 10 and 16 percent, or between 11 and 15 percent, ofthe planform area of the wing. The fuselage also has a landing gearsponson 580 for storing retractable landing gear.

As shown in FIG. 7, a rotor 520 has a blade 524 that sweeps out a circle610 during rotation that defines a rotor diameter 612. The rotor area isdefined as the area of the circle 610 swept out by the tips of rotorblades. In the case of rotors of variable diameter, the rotor area isthe maximum rotor area. The TR53 has two rotors of 2200 square feet areaeach. An aircraft that is capable of hovering has a maximum disc loadingdefined as the maximum hover weight divided by the sum of all rotorarea. The TR53 has a maximum disc loading of 27.3 square feet. Preferredaircraft have a plurality of rotors rotor sized and dimensioned to hoverthe aircraft with a rotor disc loading of at most 20, 30, 40, 50, oreven 60 pounds per square foot at a maximum hover weight.

FIG. 8 contains a table with dimensions and physical characteristics ofa preferred TR53.

Aircraft Structure, Mechanical Systems, and Weights

At a constant disc loading, the thrust produced by a rotorcraft isrelatively proportional to the rotor diameter squared while the rotorweight is proportional to the rotor diameter cubed. This is known as thesquare-cube law in the industry and results in both an undesired trendof increased disc loading in larger rotorcraft, and extreme difficultyin designing very large rotorcraft. Large rotor designs increase discloading in order to reduce the diameter appropriate for a given vehicleweight. Furthermore, prior art large rotor designs have articulated hubsystems to minimize blade flap and hub bending moments. However, thecomplexity of articulated rotor systems as found on most largehelicopters contributes significantly to their high cost, relativelyshort lifetime, and high failure rate.

Simpler rotor designs exist, known as hingeless rotors, which haveadvantages in cost, lifetime, and failure rate. Because of the highbending loads associated with hingeless rotors, and the effects of thesquare-cube law, hingeless rotors have been limited to small rotors inorder to avoid the increase in rotor weight resulting from these loads.U.S. patent application Ser. No. 12/427,961 teaches hubs for largehingeless rotors capable of withstanding high bending moments.

The TR53 is envisioned as a high composite content aircraft takingadvantage of several aspects of composite materials and manufacturingtechnology not currently integrated. The TR53 is currently contemplatedto operate at a relatively high rotor disc loading (as compared tohelicopters), of 23 pounds per square foot, with potential capacity of30 pounds per square foot, or more. Such a high disc loading combinedwith the need for lightweight structure produces a demand forlightweight and stiff rotor blades. The rotor blades of preferredembodiments are of a substantially all-composite structure, using highmodulus graphite in structural areas important to the required bladestiffness. As described in Self-Tooling Composite Structure, U.S. patentapplication Ser. No. 12/200,534, the blade structural spar isadvantageously of a multi-celled design that increases bucklingstrength, and creates a relatively strong interface between the spar capand web by interweaving plies.

The 23-30 pounds per square foot of disc loading assume a conservativestructural safety factor. Using less conservative safety factors, thesubject matter herein is enabling for disc loading of up to 40, 50, 60or even 70 pounds per square foot. Independently, the thickness ofcontemplated blade structural laminates or hub wall thicknesses could beincreased to allow for a rotor disc loading of up to 40, 50, 60 or even70 pounds per square foot.

U.S. Pat. No. 6,641,365, Optimum Speed Tilt-Rotors, and U.S. Pat. No.6,007,298, Optimum Speed Rotor provide major steps forward in rotorcraftutility, maneuverability and performance. Among other things, thesepatents teach low rotor blade weights, and appropriate structuraldynamic solutions to avoid aeroelastic instability. Significantly,however, OSR or OSTR rotors develop high blade bending moments and mastmoments, which increase with rotor disc loading. In OSR and OSTR rotors,which necessarily are relatively light and stiff, blade flap moment androtor mast moments dominate other rotor loads. The preferred embodimentsof '298 and '365, as well as the implementation of the OSR patent in theBoeing Hummingbird A160, were for low rotor disc loadings, typicallyless than 6 pounds per square foot.

Composite Blade Root Structure, U.S. patent application Ser. No.12/397,833, Self-Tooling Composite Structure, U.S. patent applicationSer. No. 12/200,534, and Wing And Blade Structure Using PultrudedComposites, U.S. patent application Ser. No. 12/397,141 providecomposite structures and manufacturing methods to achievehigh-stiffness, lightweight affordable blades which are capable ofsustaining the high loads typical of a high disc loading OSTR rotor.

High Compressive Strength Fiber-Placed Composites, U.S. PatentApplication No. 61/099,865, Structural Enclosure For An AircraftPropulsion System, U.S. patent application Ser. No. 12/254,971,Composite Bulkhead And Skin Construction, U.S. patent application Ser.No. 12/246,904, Automated Prototyping Of A Composite Airframe, U.S.patent application Ser. No. 12/396,927, High Quality Out-Of-AutoclaveComposites, U.S. Patent Application No. 61/047,877 and Rotorcraft withIntegrated Spar and Tilt Trunnion, U.S. Patent Application No.61/047,853 provide composite structural design, materials andmanufacturing processes to achieve substantial reduction in weight andreduction in manufacturing cost of composite airframe.

High Performance Outboard Section for Rotor Blades, U.S. patentapplication Ser. No. 11/505,157, Shaped Blade for Reduced Loads andVibration, U.S. patent application Ser. No. 11/505,040 and RotorcraftWith Opposing Roll Mast Moment U.S. patent application Ser. No.11/505,066 provide systems and methods related to the rotor bladesand/or rotor operations, mainly in order to improve performance andreduce loads and vibrations.

Blade section structural analysis on the TR53 has been performed usingCosmos™/M show no buckling, and a relatively even stress distribution incompression along the upper cap and tension along the bottom cap underupward bending. Even stress distributions are conducive to long bladelife and efficient structural design.

In preferred embodiments, a thin, aerodynamically efficient blade shapetransitions to a circular attachment at the hub interface as describedin Composite Blade Root Structure, U.S. patent application Ser. No.12/397,833. Additionally, the rotor blade can advantageously usepultruded high modulus composite materials in the lamination, asdescribed in Wing and Blade Structure Using Pultruded Composites, U.S.patent application Ser. No. 12/397,141, creating a structure with a highcompression strength and high overall stiffness. Laminations canconstitute approximately 10 unidirectional plies of 0.007″ thickness,interlaid with one or two biased plies of the same thickness. Two orthree unidirectional pultrusions of 0.02″-0.035″ thickness and 1″-2″width can be laminated in place of the individual unidirectional pliesto create a pultrusion based lamination. This is especially preferred inthe upper cap of the blade spar, as it will benefit most from theincreased compression strength that the pultrusions provide. Analysis,conducted with Cosmos™/M software, shows that pultrusions can be madethin enough such that their terminations within a laminate do not inducestresses that would damage the laminate. In preferred embodiments,several pultruded plies are interlaid with biased prepreg layers. Insome cases, pultrusions can also be tapered at their terminations toreduce the stress concentrations that are developed.

In preferred embodiments, the hub is capable of providing lightweightefficient blade feathering support under the large bending momentsgenerated by rotor blades. U.S. Patent Application No. 61/047,167describes a preferred integrated lubrication and cooling system disposedwithin a tailored thickness bearing support structure. The use of thisdesign creates a lightweight and efficient transfer of the large blademoment loads to the monocoque nacelle structure. The '167 patentapplication further teaches stiffness, reliability, fatigue life and lowmaintenance required for low weight and affordable cost.

The thickness distribution and shape of the hub structure of the TR53 ispreferably tailored to limit the deflection of the feather bearing.Deflection of points around the feather bearing perimeter have beenanalyzed versus distance from the center of the bearing, showing therelatively low deflection of the bearing out of the plane normal to theblade feather axis as deflected under load. Low deflection of thebearing out of the plane contributes to long bearing life. Analysis ofthe TR53 hub structure done with CATIA™ and NASTRAN™ software showsmaximum von Mises stresses within acceptable material limits for theenvisioned two-part forged titanium or steel design.

The preferred hub system of the TR53 integrates a closed lubrication andcooling system that supplies both the gearbox in the non-rotatingairframe and the feather bearings and actuators in the rotating hubframe. The TR53 is envisioned to use substantially all-electricactuation to provide additional weight savings. Furthermore, especiallypreferred embodiments include individual electric actuation for eachrotor blade, thereby providing the aircraft with individual bladecontrol, and the resultant weight, safety, and reliability benefits.

U.S. Patent Application No. 60/981,559 teaches a monocoque nacellestructure. Preferred nacelles have a large diameter and are constructedof high strength composite materials to enable the transfer ofrelatively large loads from a rotor to a wing and fuselage. In preferreddesigns, a limited portion of the nacelle structure serves as a primarymonocoque load path from a rotor hub to a tilting nacelle supportstructure or trunnion. Other structural portions of the nacellepreferably comprise lightweight sandwich construction composite panels,and can be shaped to improve efficiency in high-speed flight.

One preferred nacelle primary structure is illustrated in FIG. 9.Unidirectional fibers 950 are evenly distributed around thecircumference of the nacelle to hub interface 952 converge at severaldiscrete points near the attachment of the nacelle to the spinnionstructure 954. The nacelle is fastened to the spinnion using titaniumfittings 956. Details of this mechanical interface between the nacelleand hub bearing have been analyzed. A composite-to-metal transition thatallows the hub bearing to be fastened to the nacelle is contemplated.Thus, it is contemplated that loads generated by a TR53 rotor can beadvantageously transferred from the rotor to the nacelle and wing usinga spar structure joining the nacelle to the inboard wing.

Unlike most tilt rotor aircraft, the TR53 uses a spinnion. U.S. PatentApplication No. 61/047,853, teaches such a structure, which combines thefunction of a high moment capacity conversion spindle and an outboardwing spar, thereby reducing weight for a higher capacity structure andlowering part count.

The TR53 is designed to be capable of transitioning the position of thenacelle under heavy moment loads not seen by prior art tilt rotoraircraft. Additionally, the hingeless nature of the TR53 rotor allowsfor an advantageous method of creating a control yaw moment in thevehicle without vectoring the rotor thrust. Tilt Actuation for aRotorcraft, U.S. Patent Application No. 61/044,429, enables thisfunctionality, as well as providing a lightweight geared down mechanismfor tilting the rotor. Also envisioned in the '429 application aremethods of operating the tilt mechanism with moment loads generated bythe rotor, both to increase fault tolerance and increase performance.

Either of two gearbox configurations, demonstrated in Torque BalancingGearbox, U.S. patent application Ser. No. 12/399,594 or LightweightReduction Gearbox, U.S. Pat. No. 7,500,935, is envisioned for use in theTR53. Both enable high torque capacity operation at weightssignificantly lighter than those achieved by conventional designs.

The TR53 is contemplated to use several of the aforementioned patentpending systems and methods to produce a lighter, safer and moreefficient aircraft. In addition, the airframe is designed toaggressively reduce weight and cost. Aircraft manufacturers have begunto embrace composite materials, for example, the Beechcraft™ Starship,Raytheon™ Premier™ Boeing™ 787 each make use of carbon compositematerials in primary structure. However, these aircraft gain only asubset of the advantages achievable through use of composite materialsand manufacturing technology. In the prior art construction of abulkhead to skin assembly of a Boeing™ 787, each individual compositebulkhead is fastened to the skin through an intermediary piece, greatlyincreasing part count. The total titanium rivet count in one barrelsection of the Boeing™ 787 fuselage is over 10,000 pieces. Although theBoeing™ 787 is almost 50% composite materials, titanium represents 15%of the overall weight of the aircraft. In addition to this, the fuselageis assembled from more than 6 barrel sections.

By contrast, it is contemplated that the TR53 fuselage could beassembled in 2-3 segments as shown in FIG. 10. Each segment preferablyhas a length that is 20%, 30%, or even 40% of the length of the entireaircraft. Each of the fuselage segments 1010 could be built up in acomposite tool manufactured from low-temperature tooling materials,including for example the process described in U.S. patent applicationSer. No. 12/396,927. These fuselage segments 1010 would be laminatedwith an automated fiber placement, using out-of-autoclave resin systems.By using out-of-autoclave low-temp-high-temp two-stage resins, the laborand capital expenses of the autoclave are eliminated. These hightemperature capable tools have the locating features necessary to doubleas assembly tools during the bonding of composite bulkheads into thefuselage sections further reducing the step of transferring the partfrom one tool to another, and reducing the overall tool count.

As described in Composite Bulkhead and Skin Construction, U.S. patentapplication Ser. No. 12/246,904, the bulkhead frames can use pultrudedcarbon materials in the inner caps, and the outer cap material can belaminated into the outer skin during the automated fiber placementprocess. The skin of the fuselage can be a highly orthotropic laminatesupported by co-cured laminated hat stringers and bulkheads with cutoutsfor stringers, or it can be a hybrid honeycomb-cored sandwich skin andbulkhead supported structure. Ideally, the tension in the fibers can becontrolled during fiber placement to decrease the waviness of thelaminate, and thus increase the compressive strength of the laminatewhere needed. Aspects of preferred systems and methods are disclosed inU.S. Patent Application No. 60/979,630.

Several innovative design features combine to substantially reduce theweight of the contemplated TR53 airframe relative to prior artrotorcraft. In FIG. 10, the wing box 1030 is made from a single piecethat fits between the left nacelle 1050 and right nacelle (not shown).The attachment of the wing to the fuselage is through a pinnedattachment that allows the wing to bend in flap. Thus, bending loads inthe wing are transferred to the fuselage primarily as shear, notbending. Replaceable components such as wing flaperons 1032, 1044 andslats 1042 are made of lightweight composite sandwich construction.Other components include a horizontal tail 1064 and vertical tail 1062,and fuselage bulkheads 1090.

Wing and Blade Structure Using Pultruded Composites, U.S. patentapplication Ser. No. 12/397,141, also teaches the concept of usingpultrusion based laminates in the wing box of an aircraft. Currentlycontemplated versions of the TR53 integrate this concept into theinboard wing box, rotor blade, and spinnion. Modal and structuralanalysis performed on this wing structure with Cosmos™/M and spreadsheettools have verified the design.

FIG. 11 is a table containing preferred weights including preferredempty weights and preferred maximum hover weights (design gross weights)as calculated for two versions of the TR53 as verified using theanalyses described above. This table contains details of thecalculations showing the empty weight fraction of a civil passengerversion of the TR53 to be 0.55 and the empty weight fraction of a civilcargo version of the TR53 to be 0.34. Preferred aircraft have an emptyweight is at most 58%, 60%, or 65% of the aircraft maximum hover weight.

Aerodynamic Efficiency

The TR53 is the first VTOL transport aircraft contemplated torealistically compete with traditional fixed-wing jet and propellertransports. To achieve commercial viability in a tough market dominatedby jet transports, the TR53 combines a helicopter-like disc loading (ahigher disc loading requires higher installed power for hover takeoffand landing) with high hover figure-of-merit and aerodynamic performanceneeded for efficient cruise.

Aerodynamic Efficiency—Blade Root and Spinner for High Speed Cruise

The TR53 is designed to cruise efficiently at Mach 0.65. This is about35% faster that the prior art Bell™ V-22 tiltrotor and 30% faster thanthe prior art Bell™ BA609 tilt-rotor. This cruise speed is importantboth for aircraft passenger transport productivity goals in terms ofpassenger miles per day and for the important goal of reduceddoor-to-door travel time. To achieve these goals, the present embodimentcontemplates shaping the inboard blade planform to increase chord whilereducing thickness ratio.

In preferred embodiments, this is combined with judicious design andimprovement of blade root airfoils, which are depicted in FIG. 12. Ablade 1200 comprises a leading edge 1204 and trailing edge 1202 arrangedaround a pitch axis 1206. A preferred embodiment root airfoil section1210 has a thickness ratio, airfoil thickness 1212 divided by airfoilchord 1214, t/c, of only 22.5%. The thickness ratios diminish withincreasing blade span as shown in FIG. 12, such that an airfoil 1290 at90% span has a thickness ratio of 11.4%. Achieving a thickness ratio of22.5% at the blade root while ensuring that the desired structuraldynamic characteristics (and the desired low blade weight per blade flapmoment capability) are achieved, such that the first blade flap and lagfrequencies are greater than 2/rev at the maximum hover rpm, requiresspecial blade composite construction methods, described previously. Theterm “2/rev” refers to a frequency associated with a speed twice that ofthe rotor rotational revolution, and would mark an event happening attwice per rotor revolution.

Lift-to-drag polars for preferred embodiment rotor airfoils show reduceddrag even at section Mach numbers of 0.65 to 0.70. Maintaining low draglevels at high Mach number is traditionally difficult. For example, itis difficult to maintain a sectional drag coefficient of 0.045 at Mach0.70 using prior art airfoil sections having a thickness ratio of 26.5%or more. For the TR53, a lift-to-drag polar for an especially preferredembodiment root section airfoil having a thickness ratio of 22.5% wascalculated that shows that a drag coefficient of only 0.025 could beobtained a preferred root airfoil section (see FIG. 12), a majorimprovement over the prior art. The lift-to-drag polars for thinnersections further outboard on the blade have thickness ratios of 20% and11.4% and 10% at the tip. These preferred embodiment airfoil designsprovide very low drag and a broad operating envelope. These lift-to-dragpolars were obtained using computa-tional fluid dynamics software, MSES,a coupled viscous and inviscid Euler approach well known in theindustry, and ANSYS™ Fluent.

The TR53 also has a rotor spinner fairing. A spinner creates blockagefor the inboard blade sections, further raising the local flow Machnumber. Computational fluid dynamics simulations have analyzed theairflow around blade roots, spinner, and wing at cruise speeds includingMach 0.65. As shown in FIG. 13, a dramatically area-ruled spinner iscontemplated that can lower the Mach numbers by 0.01-0.03 at rotor bladesections inboard of 25% of span, depending on the degree of area rulingand spanwise position of the section, substantially increasing theenvelope for efficient cruise. In the TR53, shaping the tiltrotorspinner is calculated to increase the efficient cruise Mach number ofthe rotor by about 0.02.

FIG. 13 is a side-view illustration of a preferred blade-nacelleinterface 1300. A blade comprises a spar 1310, aerodynamic fairing 1312,section of the spar near the shank 1320, and cuff 1330 acting as atransition to a blade shank 1340. The blade shank 1340 is coupled to arotating hub 1350, which in turn is coupled to a tilting nacelle 1370through a rotating interface 1372. The hub 1350 and blade shank 1340 areat least partly disposed in a spinner 1360, which acts as an aerodynamicfairing. The spinner is advantageously configured with a concave region1362, reducing the spinner diameter by 3%, 5%, or even 10% from amaximum spinner diameter. This concave region 1362 effectively slows theairflow as it passes over the thick blade root region, allowing theaircraft to fly faster. The blade cuff 1330 and aerodynamic fairing 1312are also configured to provide smooth airflow into an engine inlet 1364.

Application of the teachings found in this specification allow forrelatively smooth airflow at Mach 0.55, 0.60, or even 0.65 around thenacelle 1370 and the absence of a shock wave at the engine inlet 1360 tothe engine even given the relatively thick blade aerodynamic fairing1312 and associated airfoils. As used herein, a “relatively thick”airfoil has a thickness-to-chord ratio of at least 18%, 20%, 23%, 27%,or even 30%.

In preferred embodiments, the hub 1350 is coupled to a reduction gearbox1384, two-speed shifting gearbox 1382, and turboshaft turbine engine1380. Along with the rotor and airframe design, the configuration of thereduction gearbox 1384, shifting gearbox 1382, and turboshaft engine1380 allow the aircraft to sustain cruise flight with the rotoroperating at a rotational speed no greater than 40%, 50%, 60%, 75%, or80% of its operational maximum rotational speed. The turboshaft engine1380 is advantageously sized and dimensioned both to power the rotors tolift the aircraft vertically, and to provide sufficient power toovercome an aerodynamic drag of the aircraft in sustained high-speedforward flight at Mach numbers of 0.5, 0.55, 0.6, or even 0.65. As usedherein, the term “sustained cruise flight” means a steady, level flightcondition (excludes maneuvering flight) substantially in equilibrium fora continuous period of at least 10 minutes. For example, an aircraft ona flight from San Francisco to Los Angeles might take off vertically,convert to airplane mode, climb to a designated altitude of 31000 feet,engage in sustained cruise flight at that altitude for 45 minutes, thendescend and vertically land.

Aerodynamic design is also important for proper operation of a turbineengine. In preferred embodiments, for example, the engine air intake, isplaced behind the plane of the rotor. Because of this, the inlet ingeststhe rotor wake flow. Still further, the blade root design, spinnershape, and inlet design of preferred embodiments are all iterated in acoupled environment that tries to ensure that the flow entering theengine is in such a state that engine limits for safe and efficientoperation are satisfied. Computational fluid dynamics simulation resultsshow the trailed blade wake in front of the engine inlet can beaccommodated without undue loss of efficiency. The inlet of the TR53 isdesigned to provide high efficiency in hover while providing a largemass flow of air to the engine. The inlet of the TR53 is also designedto provide high efficiency in cruise while providing a reduced mass flowof air to the engine and low spillage drag when the aircraft issimulated to be cruising at a speed of Mach 0.65.

The aforementioned design process and aerodynamic design of the rotorand inlet for the TR53 also assists in maintaining sufficient engineoperational margins during flight.

Aerodynamic Efficiency—Efficient Transonic Wing

A large high-aspect ratio wing can assist in achieving cruiseefficiency, and in closing the productivity gaps with traditionalfixed-wing transports. Currently contemplated versions of the TR53include a wing capable of supporting tilting structure, that issufficiently stiff to remain free of aeroelastic flutter to Mach 0.65and beyond, and that also is sufficiently thin to be capable ofefficient flight at Mach 0.65. These are usually conflictingrequirements, but the present inventive subject matter provides aresolution of the conflict by use of a judiciously-designed compositewing structure and special airfoil designs. It should be noted thatlightweight rigid rotors, such as those taught in OSTR, U.S. Pat. No.6,641,365 can delay the onset of whirl flutter to a higher flight speedthan a conventional rotor system. A preferred wing is shown in FIG. 5 asinboard wing section 502 along with outboard wing sections 508, 509.

Prior art tiltrotor wing designs, the Bell™ V-22 and Bell™ BA609, werethick (23% thickness ratio) and of low aspect ratio, 5.5, to maintain alower weight while remaining aero-structurally stable to Mach 0.45. As aresult, these prior art designs were unable to achieve high cruiseefficiency.

By contrast, highly efficient transonic wings for a transport aircraft,such as that disclosed herein, have high aspect ratio and low thicknessratio, with an appropriate degree of wing sweep. A currently mostpreferred wing for the TR53 has a very high wing aspect ratio of 16.3and a low thickness ratio of 20.4%, and an inboard wing sweep of 5°.This provides unprecedented cruise efficiency for a tilt-rotor. Also,the inboard wing for the TR53 is currently contemplated to have a taperratio of 0.86 to increase bending stiffness within acceptable weight.Teachings on improving wing efficiency in tilt-rotor aircraft can befound in U.S. patent application Ser. Nos. 11/505,067 and 11/505,025.

The TR53 is also contemplated to include airfoils that increase themaximum lift limit at a high cruise speed. This is known as the M²C_(L)limit in the industry, and results from the product of the square of thecruise Mach and the lift coefficient. These airfoils provide most of thelift on the leading edge under camber and trailing edge under camber,while providing a thick mid-section for the required wing stiffness inbending and torsion. Special care is also given to increasing airfoilsection thickness at the chordwise location of the nacelle tiltspinnion. FIG. 14A depicts a preferred wing airfoil 1402 that is 23%thick with an undeflected cruise flap 1404 and partially deflected flap1406. FIG. 14B depicts an especially preferred 20.4% thick inboard wingairfoil 1412 and undeflected flap 1414 section. The airfoil has a chord1416 and thickness 1418. The thickness ratio, expressed as a percentage,is computed by dividing the maximum thickness 1418 by the chord 1416with any flaps 1414 or slats in the undeflected configuration.Computational fluid dynamics results on this section at Mach 0.65 show alarge laminar drag bucket, with a usable operating regime having asectional drag coefficient of 0.006 at Mach 0.65, a significantimprovement over the prior art. The outboard wing section was equippedwith a novel reflexed flap deployed to minimize separation and drag.Computational fluid dynamics results generated using ANSYS™ Fluent at acruise Mach 0.65 show good overall wing pressure distribution andefficiency of a preferred embodiment.

The high aspect ratio wing in preferred embodiments creates lessblockage area and download force (an aerodynamic force produced by therotor on a wing or fuselage that opposes the thrust of the rotor) inhover. Prior art tiltrotor designs, the Bell™ V-22 and Bell™ BA609, haddownload of 11% and 9%, respectively. The TR53 includes a trailing edgeflap of 40% of the wing section chord, which is calculated to provide atotal download of 4.7% using the computational fluid dynamics codeANSYS™ Fluent. This trailing edge flap is further optimized to reducedownload. For comparison with the prior art, a download reduction from11% to 4.7% translates to 8% reduction in required installed power, asillustrated in FIG. 15. FIG. 15 is a graph with the hover power requiredon the vertical axis 1502 and download percentage on the horizontal axis1504. The calculated relation 1506 shows that the TR53 downloadreduction 1520 yields an 8% power reduction 1510.

The TR53 uses a tilting outboard wing as taught in U.S. patentapplication Ser. No. 11/505,025, which creates almost no download on theoutboard wing during hover and further increases the overall wing aspectratio. The outboard wing airfoils are optimized for efficient cruise atacceptable thickness ratio, and also to provide the low aerodynamicbuffet level during conversion flight from helicopter mode to wing borneflight, by using leading edge slats and trailing edge flaps. This designimprovement is useful because buffet can be objectionable to passengersas well a source of fatiguing loads for the nacelle system and wing andfuselage structures.

The high aspect ratio wing used to provide efficient cruise flight inpreferred embodiments accelerates flow on the top surface of the wing,creating a differential velocity, or wash, field in front of the wing.This accelerated flow over the top of the wing creates higher Machnumbers for prop-rotor sections rotating through the top-half of arotation, making the airfoil design problem even more important on theprop. Thus, method aspects of the present inventive subject matter arecontemplated wherein the wing geometrical design and rotor bladegeometric design are modified using the aid of computational fluiddynamics (CFD) simulations including the presence of both the wing androtating rotor. Such CFD simulation shows that sustained flight of theTR53 at speeds of Mach 0.65 and wing loadings of 60, 80, 100, or 120pounds per square foot is viable.

Still further, it is contemplated that the relatively large wing size ofthe TR53 supports cruise flight at 41,000 feet. This cruise altitude is16,000 feet higher than the prior art Bell™ V-22 and Bell™ BA609 as wellas most turboprop aircraft. Such a relatively large wing size enablesthe aircraft to cruise higher and faster while maintaining high engineoperating efficiency by substantially increasing the configurationlift-to-drag ratio.

Aerodynamic Efficiency—Low Drag Fuselage

The fuselage drag of the TR53 is an important consideration both forachieving high speed and high efficiency.

As compared to the smaller prior art Bell™ BA609 scaled to the samefuselage diameter, the TR53 is currently contemplated to have a smoothlyshaped nose that reduces drag, as shown in FIG. 6.

Additionally, the shaping of the aft section of the fuselage, oftencalled the boat tail, is important. In the TR53, the aft section is notshaped like most prior art rotorcraft, tilt-rotors, or other verticaltakeoff aircraft. Instead, the TR53 aft section has a shape similar tothat of transport aircraft designed to cruise at Mach 0.8 including forexample the Airbus™ A320. FIG. 6 shows the gentle upsweep of thefuselage aft section, with an average angle of about 14° as measured atthe lower outside mold line of the fuselage.

While a high wing configuration provides a lower drag and higherlift-to-drag ratio contribution than a low wing configuration, the needto enclose, or fair, the main landing gear adds a fairing drag. The TR53combines a unique low drag main landing gear supporting structure,landing gear doors and a stub wing-like fairing. FIG. 16 shows apreferred main landing gear supporting structure and landing gear doors1606, 1608 retracting into the stub wing-like aerodynamic fairing 1604that is coupled to a fuselage 1602. The extended landing gear 1610 andarm with shock absorbers 1612 retract into a retracted position 1620.

An area-ruled wing-fuselage fairing is used to reduce the aerodynamicdrag arising from the junction of the wing and fuselage. Embodiments(not shown) are also contemplated that increase the wing root chord toeffectively reduce the wing thickness ratio at the fuselage junction,which can alleviate high Mach effects.

The TR53 fuselage has a fineness ratio, defined as the ratio of thelength of a fuselage to its maximum diameter, of approximately 8. Thisis much higher than prior art rotorcraft. The long, slender nature ofthe fuselage can be seen in FIG. 8. Preferred embodiments have finenessratios of 6, 8, and even 10; these fineness ratios assists in providingrelatively low drag beneficial for a high utility transport aircraft.Because prior art rotorcraft were relatively slow (cruise below Mach0.45 and more typically Mach 0.15), there was little or no motivation tohave a slender low drag fuselage, and certainly no appreciation of aneed to have a high fuselage fineness ratio.

Aerodynamic Efficiency—Empennage Drag Reduction and Flight Control

Most prior art aircraft, both fixed-wing and rotary-wing, utilize anempennage, composed of aerodynamic tail surfaces, to maintainlongitudinal and directional stability and flight control. However, asempennage area increases so does the aircraft drag. In the TR53, thetotal empennage surface area is only about 18% of the wing area. This isa significant improvement as compared to the tiltrotor prior art, forexample the Bell™ V-22 tiltrotor, which has a total empennage wettedarea that is about 105% of the wing area. Fixed-wing prior art usuallyalso have higher empennage surface area to wing area ratios, asexemplified by the Bombardier™ Dash 8 Q400 at about 48% and the Piaggio™Avanti at about 67%.

This dramatic reduction in empennage area for the TR53 stems from theaircraft's use of mast moment to assist in pitch stabilization andcontrol and in directional stabilization and control of the aircraft,the latter aided by different nacelle tilt as necessary as described inU.S. Patent Application No. 61/044,429. Additional discussion of theenabling inventions that provide for this small tail area has beenprovided above in the discussion of aircraft structure and empty weight.This reduces the zero-lift drag of the TR53 by about 1% as compared tothe 48% empennage-to-wing area ratio of the Bombardier™ Dash 8 Q400, andreduces the zero-lift drag of the TR53 by 3% as compared to 105%empennage-to-wing area ratio of the Bell™ V-22. The TR53 tail surfacesare sized to provide pitch trim in cruise flight without use of rotormast moment and to provide yaw trim in the case that one rotor isfeathered. The TR53 tail surfaces also provide a marginally stableaircraft configuration in pitch and yaw. While low static stability isof limited importance for an aircraft with very high control power inpitch and yaw (such as the TR53 when using rotor control moments), theavoidance of the large yaw stability of conventional aircraft greatlyreduces the need to use rotor controls to fight the tendency to yaw intothe wind (known as weather-cocking) of conventional tiltrotor aircraftwhen landing or taking off in strong cross winds.

Further, the TR53 is contemplated to use an advanced flight controlsystem. U.S. Pat. No. 6,584,383 and U.S. patent application Ser. No.11/506,571 teach systems and methods that can be used to improveaircraft flight safety and security and reduce of flight crew workloadand training.

Method aspects of the preferred embodiments are contemplated whereby theaircraft can maintain pitch and directional control as well as maintainstability with the loss of some or all of the empennage throughapplication of mast moments generated by one or more rotors. In theTR53, this feature is contemplated to be augmented with systems fromU.S. patent application Ser. No. 11/506,044, which allows continuationof wing-born flight when one rotor or its drive system fails, and withsystems from U.S. patent application Ser. No. 11/473,979. Takentogether, these systems can provide a rotor fail-operation capability,high-redundancy roll control and tail redundancy in pitch and yaw, andtherefore continued safe operation in many contingency scenarios.Without these features, it would be difficult for a tilt rotor aircraftcould achieve the safety level required for large-scale commercial airtransportation.

Aerodynamic Efficiency—Total Drag and Lift-to-Drag Ratio

The TR53 was developed with extensive use of computational fluiddynamics (CFD) solution approaches, including the aforementioned ANSYS™Fluent, CDI™ CHARM, Overflow, and MSES tools. The use of these toolsallows the prediction and verification of the performance of the presentinventive subject matter. Further, in preferred embodiments, these toolsare used for detailed configuration and aerodynamic shaping andoptimization.

The aerodynamically efficient configuration of the TR53 is depicted inFIGS. 5, 6, and 7. Based on the use of the aforementioned simulationtools, the aerodynamic performance is computed; which shows a preferredlift-to-drag ratio of about 20 assuming turbulent free stream airconditions and thus a turbulent boundary layer, and a lift-to-drag ratioof 23 assuming favorable free-stream air conditions and extensivenatural laminar flow when the aircraft is flying at a Mach numberbetween 0.50 and 0.60. This represents a major step beyond theoperational prior art of the Bell™ V-22 and Bell™ BA609, which areestimated to have peak lift-to-drag ratios below 8 at Mach numbers ofaround 0.45.

In software simulation of TR53 cruise flight at altitudes of35,000-41,000 feet and Mach numbers between 0.60 and 0.65, alift-to-drag ratio of about 18 is calculated when conservativelyassuming a turbulent boundary layer. This result demonstrates theinherent efficiency advantage of flying the TR53 at Mach 0.65 with astraight wing (relatively increased wing span) as opposed to flying at aMach number of 0.80 with swept wing (and relatively reduced wing spanfor which configuration one of ordinary skill in the art would expect areduced lift-to-drag ratio).

Preferred aircraft are advantageously equipped with a combination oftwo, three or more features that allows the aircraft to achievesustained cruise flight at a speed greater than Mach 0.5, 0.55, or even0.6. These features include the rotor having a rotor blade with aquarter-span thickness ratio of less than 20%, 22%, 24%, or 26%, thewing airfoil with a thickness ratio less than 19%, 20%, 21%, 22%, or23%, but preferably between 19% and 22%, a nacelle having a spinner witha concave region with a diameter that is reduced by 5% or more from amaximum spinner diameter, and a turbine engine sized and dimensioned toprovide sufficient power to overcome an aerodynamic drag of the aircraftat that speed. A quarter-span thickness ratio is the rotor blade 25%radial station airfoil thickness-to-chord ratio.

Embodiments are contemplated in which a tiltrotor has a more swept wingthan the TR53. It is well known that sweeping a wing increases the highMach performance of that wing. In tilt-rotors, it is desirable to havethe wing quarter-chord and the helicopter-mode rotor rotation center bein line with the vehicle center of gravity. To accomplish this in a morehighly swept wing, the wing could be swept in a W-shape; sweeping theinboard portion of the inboard wing backwards, then reversing the sweepangle of the outboard portion of the inboard wing. Alternatively, thewing could be swept first forward then backwards. Making a tiltrotorwing in a W-configuration will require additional section thickness andwill add complexity and cost. Thus, wing sweep is not the only factorlimiting aircraft forward speed. With a low disc loading rotor withthick blades and transonic effects at the blade roots, achievingefficient cruise Mach numbers much higher than 0.70 is a challenge. Insuch an embodiment, a wing segment could have a leading edge sweep angleof 15°, 20°, or even 35°.

There is an additional positive interference effect of the TR53'srotating propellers that further decreases the drag due to lift, knownas induced drag. Performance estimates of the TR53, includingcalculation of the vehicle lift-to-drag ratio, are based oncomputational fluid dynamics computations, including the effects of theturning rotors.

It is interesting to compare a TR53 with a Boeing™ 737-600 jettransport. These two aircraft have fuselages of comparable size, andcould carry similar payloads. It is contemplated that the TR53 will beso efficient that it could be commercially competitive with the Boeing™737-600 on many routes.

Propulsion Efficiency

In general, it is difficult to make a single rotor operate efficientlyin both hover mode and cruise mode. In the TR53 this is accomplished inlarge measure by altering the speed of the rotor.

The TR53 is designed to achieve a thrust ratio of approximately 35:1between low-altitude hover flight (where 175,000 lb of vertical lift isrequired for helicopter-mode maneuver including blade stall margin) andlow altitude cruise flight (where only 5000 pounds of thrust isrequired), and possibly 10:1 in other flight conditions. This isaccomplished by varying the rotor speed sufficiently to maintain arelatively high blade loading (distributed sectional lift coefficient)and rotor efficiency η_(p) in cruise mode.

An aircraft that has rotors sized and positioned to lift the aircraftvertically means that the aircraft is capable of substantially verticaltakeoff at sea level using thrust generated by spinning rotors. Inpractice, such rotors must be of adequate diameter to provide sufficientlifting force.

An aircraft having a rotor has an operational maximum rotational speed(usually given in rpm) associated with that rotor. As used herein theterm “operational maximum rotational speed” is the maximum steady speedat which a rotor can safely rotate during takeoff. In practice, theoperational maximum rotational speed of a rotor might be set bystructural integrity limits, vibration limits, power limits, stabilityor dynamic limits, gearbox limits, drivetrain limits, drag limits, noiselimits, or some other concern. Realistic aircraft typically have rotoroperation limits noted on cockpit instruments, pilot controls, flightmanuals, design documents, or elsewhere. For a version of the TR53 witha Rolls-Royce™ AE1107C engine, the operational maximum rotational speedis calculated to be about 285 rpm. In contemplated hovering flight, theTR53 could fly with substantial payloads at rotational speeds of 250,270, or 285 rpm. In contemplated cruise flight, the TR53 could fly witha rotor speed of 130, 145, or 160 rpm. Naturally, these rotationalspeeds are merely examples, and many other rotational speeds might beachieved depending on the flight condition.

The preferred rotor and drive train for the TR53 is designed to operateat 49% of maximum rpm for efficient cruise (long-range flight) and downto 26% of maximum rpm for loiter (long endurance flight). As an example,a TR53 might cruise at high altitude with a reduced rotational speed of44% of the operational maximum rpm, calculated to correspond to 50% ofthe maximum lift coefficient of a working section of a rotor blade,which provides good propeller efficiency. Rotor propulsive efficiency isshown in FIG. 17, which is a graph of the result of a computationalfluid dynamics simulation of a preferred rotor, where the horizontalaxis 1702 is the propeller advance ratio and the vertical axis 1704 isthe propeller power coefficient. FIG. 17 is a calculated rotorefficiency map at Mach 0.60 assuming fully turbulent flow, showing thata large plateau 1706 of rotor efficiency above 84% is available. The useof variable rotor speed allows the operator to select values of advanceratio and power coefficient to continually stay inside this plateau 1706over a wide range of flight speeds and altitudes. Further, the rotor iscalculated to be capable of efficient operation at higher Mach numbersas well, here at up to 84% efficient at Mach 0.65 with some laminarflow.

In summary, a rotor system can advantageously vary its rotational speedto maintain a relatively high propulsive efficiency. However, knownturboshaft engines are unable to maintain high efficiency over anequivalent rotor speed range. For this reason, the TR53 utilizes a2-speed transmission with 1.75:1 gear ratio to maintain a low rotorspeed but a relatively higher engine speed. Aspects of preferredspeed-changing systems and methods are described in U.S. patentapplication Ser. Nos. 11/473,978 and 12/399,291. Other aspects ofpreferred drive trains and transmissions are described in U.S. patentapplication Ser. Nos. 11/473,969 and 12/399,594.

Preferred aircraft are advantageously equipped with an engine and drivetrain capable of operating the rotor at a cruise rpm of 70% to 75%, 50%to 60%, or even 30% to 50% of the maximum rpm in hover. It is known thatturboshaft engines can operate over some range of output rotationalspeeds, for example, 60% to 100% of an engine maximum rotational speed.The extraordinarily wide output speed range of preferred rotors anddrive trains implies that the engine either be able to operate over anextraordinarily wide speed band (of which known prior art turboshaftengines are not capable) or that the aircraft be equipped with amultiple-speed ratio gearbox.

One preferred engine for the TR53 is the Rolls-Royce™ AE1107C. However,since it any suitable engine or engines could be used to power thecontemplated rotorcraft, the discus-sions of Specific Fuel Consumption(SFC) herein is made with respect to both current production engines aswell as advanced engines. Current turboshaft engines lag in performancebehind their large turbofan counterparts. The highest performanceturboshaft engine currently entering production is the European EPI™TP400-DG, which provides 11,000 horsepower at sea level and estimatedcruise SFC of 0.36 pounds of fuel per horsepower-hour. With this SFCfigure, the TR53 is calculated to provide an estimated thrust SFC of0.44 pounds of fuel per pound thrust-hour at a cruise speed of Mach0.65. Using modern turbofan technology, the TR53 is calculated toprovide a thrust SFC of 0.40 pounds of fuel per pound thrust-hour at acruise speed of Mach 0.65.

Integrated Cruise Efficiency

Achieving a high cruise performance in a tiltrotor is a formidablechallenge. In order to hover the vehicle out of ground effect andmaintain a degree of hover capability with one engine inoperative, thevehicle must have a large total installed power. However, if the vehicleis aerodynamically efficient with a high lift-to-drag ratio, a smallpercentage of the total installed power is needed for cruise. Varyingthe rotor speed is a step towards overcoming this gap, but the increasedrotor driving torque must be accommodated in the vehicle and drivetraindesign. Further, in turbine engines, varying the engine outputrotational speed tends to reduce the engine efficiency and increase theengine specific fuel consumption, driving down overall cruiseefficiency. Lower air densities at increasing cruise altitudes can beleveraged to offset this effect. However, in order to cruise at a higheraltitude, a larger wing and high M²C_(L) limit are needed. As has beendescribed herein, the TR53 creatively combines design elements toachieve efficient cruise performance and a robust hover out of groundeffect capability.

For a TR53 cruising at Mach 0.65 above 36,000 feet (372.7 knots) at aweight of 100,000 pounds and at L/D=20, the drag is estimated to be5,000 pounds, and power required for cruise is calculated to be 5,723horsepower. With a thrust SFC of 0.4 pounds of fuel per poundthrust-hour, the fuel consumption would therefore be 2,289 pounds perhour or 6.14 pounds per nautical mile.

For units more popular in ground transportation, with a fuel density of6.8 pounds per gallon and converting units into statute miles, the TR53is calculated to cruise at 0.78 gallons per mile or 1.28 miles pergallon with 120 seats. Compared to a family car with 5 seats, this wouldbe equivalent to 29 miles per gallon per 5 seats at 430 miles per hourand 36,000 to 41,000 feet.

It is contemplated that a TR53 cargo version would have a loweroperating empty weight (estimated at 43,000 pounds) because thestrengthened floor for carrying cargo is substantially lighter than thewindows, seats and other amenities for carrying passengers.

A contemplated civil cargo version would be capable of carrying up to42,000 pounds of cargo to a range of 5,000 nautical miles. That vehicleis estimated to have a normalized cruise efficiency (tons of payloadtimes nautical miles of range divided by pounds of fuel burned dividedby engine specific fuel consumption expressed in pounds of fuel burnedper horsepower per hour) of 9.87 with 42,000 pounds of payload whenflown to an average of 2,000 nautical miles. That represents a largeimprovement over the prior art.

Preferred aircraft have a maximum range of at least 1000, 1500, 2000,2500, or even 3000 miles carrying its maximum payload. Viewed from analternative perspective, preferred aircraft have a fuel capacitysufficient to fly at least 1000 miles, carrying at least 40%, 50%, 70%,or even 100% of its maximum payload, without refueling and after takingoff vertically. As used herein, the term maximum range means the maximumdistance an aircraft can travel without refueling after taking offvertically.

Vibration, Dynamics, and Noise

The TR53 is designed to achieve similar levels of comfort in terms ofnoise and vibration as conventional fixed-wing turboprop aircraft.

Method aspects of the present inventive subject matter are contemplatedthat allow for vibration reduction. For the TR53, this is achieved byusing the primary flight control system for vibratory load alleviationin different flight modes. In wing-borne forward flight of the preferredembodiment, cyclic pitch control is used to counteract the effects ofthe wing wash field on the turning rotor. Higher-harmonics of bladepitch can also be used, as the TR53 is designed with provision forindividual blade control. Similar methods are applied in rotor-borneflight of preferred embodiments. Additionally, noise levels are reducedwith a combination of blade tip shaping, lowered rotor rpm and reducedthickness blade sections, especially in the outboard sections of theblade.

For the TR53, it is contemplated that interior noise could be reduced byusing a high aspect ratio inner wing that places the rotor further awayfrom the fuselage than the minimum geometric constraint, as shown inFIG. 7.

As part of the design process of the TR53, the structure is analyzed tobe free of aeroelastic instabilities and other structural dynamicproblems. This analysis is performed using CAMRAD™ II software. Theinputs to this analysis and design process used for the preferredembodiments are the details of the aircraft structure and geometric andaerodynamic characteristics as described above. Using this analysis anddesign approach, it was determined that TR53 has a wing design and rotordesign sufficient to remain free of whirl flutter throughout the flightenvelope.

Design for Growth

For the TR53, the rotor is of 53 foot diameter, while the wing, hub,gearbox, and actuators are advantageously designed to have the capacityto accommodate a 56 foot diameter rotor. When the airframe is equippedwith a 53 foot rotor, this additional spacing between the loaded bladesand passenger fuselage reduces noise and annoyance for passengers.However, it is envisioned that the TR53 could be upgraded with a larger56 foot diameter rotor that would allow for expanded aircraft capabilityand performance.

In an especially preferred embodiment, the rotors of the TR53 can beupgraded from a 53 foot diameter to a 56 foot diameter by simplychanging the blades. This is due to the relative over-design on the hub,actuators, gearbox, and nacelle. Thus, the TR53 design works well with7,000 horsepower per engine as describe herein, but the design is alsocalculated to work well with 12,000 horsepower per engine, with thelarger engines being either of current or advanced technology. The TR53fuselage has a cross section similar in size to that of the prior artAirbus™ A318, preferably equipped to transport six passengers abreast.This fuselage could be stretched to accommodate 210 passengers or more,in a similar manner to the stretch of the prior art Airbus™ A318 to theprior art Airbus™ A321 fixed-wing jet transports. Use of OSTR technologyallows for stretching a vertical takeoff transport aircraft because highmast moment capability is needed to accommodate aircraftcenter-of-gravity offsets and to precisely maintain fuselage attitude.Finally, the use of OSTR rotors and a multiple-speed transmissioncreates unique growth options. In general, the transmission torque limitfor such a vehicle is set by the cruise condition, when the outputrotational speed is low and the torque is high. In hover, the powerrequirement is high, but the torque is lower because of the higheroutput rotational speed. Thus, additional engine power can be easilyaccommodated in the aircraft.

The same principles that provide the TR53 with high efficiency in hoverand forward flight could also be applied on a larger aircraft. Forexample, a tiltrotor that is approximately 1.414 times the linear scaleof the TR53 would result in an aircraft having 75-foot diameter rotors.Approximately twice the installed power would be needed, which could beaccommodated by doubling the number of engines.

Medium Modulus Graphite Fibers

A new preferred embodiment is described below, in which the rotor bladeis redesigned to enable the use of a high strength/medium modulusgraphite fiber in place of the high modulus graphite fiber used beforefor OSR and OSTR blades. Here, high modulus graphite refers to carbonfiber with a tensile modulus of 70-80 million pounds per square inch,characteristic of fibers such as Toray™ M55J™. In contrast, highstrength/medium modulus graphite refers to a carbon fiber with a tensilemodulus of 40-50 million pounds per square inch, characteristic of butnot limited to fibers such as Toray™ T800™ and T1000™, and Hexcel™ IM7™,IM8™, and IM10™.

The use of a high strength/medium modulus fiber offers many advantagesover the previous embodiment using high modulus fiber. First, mediummodulus fiber offers an elongation at failure range of 1.8% to 2.2%,compared to 0.75% for high modulus fiber; this means structure usingmedium modulus fiber will be less brittle and more damage tolerant.Second, medium modulus fiber offers not only more elongation at failurebut higher tensile and compression strength, allowing for stronger,lighter structure in strength-critical areas. Finally, medium modulusfiber is widely used in aircraft such as the Boeing 787™ and AirbusA350™; as a result, it is much easier to source and often half the costof high modulus fiber such as M55J.

FIGS. 18, 19A-19F and 20 are top, side and front views, to scale, of anespecially preferred rotor blade of the current invention showing thesubstantially widened blade root chord (9.3% of rotor diameter ascompared to 5.9% in U.S. Pat. No. 6,641,365).

FIG. 21 is a table similar to FIG. 5A of U.S. Pat. No. 6,641,365,demonstrating what material properties would be needed to achieve thebending stiffness and section weights listed in the table for the givenblade and effective spar dimensions.

The numbers were calculated by first calculating the composite capthickness that would be needed to achieve the given flap bendingstiffness for two different types of fiber: a medium modulus fiber(assumed modulus of 44 msi) and a high modulus fiber (assumed modulus of78 msi). This calculation relied on four assumptions: 1) the cap is thinrelative to the effective spar depth, 2) the entire spar width is at themaximum given spar depth, 3) 90% of the fibers in the spar cap areoriented along the blade radius, and 4) all of the composite materialused in the cap is evenly distributed between the upper and lower cap.All four of these assumptions will tend to overestimate the bendingstiffness of the blade, meaning that the calculated thickness can beconsidered the minimum required cap thickness to achieve the statedbending stiffness.

Then, using an industry-accepted density for carbon fiber compositematerial (which is generally close to being the same for most fibertypes, Toray™ M55J™ fibers are 6% higher density than Toray™ T800™fibers), the minimum required cap thickness was used to determine theminimum weight of the combined upper and lower spar cap laminates. Thiswas subsequently divided by the total blade sectional weight given inU.S. Pat. No. 6,641,365 to determine if the combined stiffness, totalblade weight, and geometry could reasonably achieved by each type offiber.

A medium modulus fiber would require a sectional spar cap weight equalto at least 120% of the total blade sectional weight, which is clearlyimpossible, while a high modulus fiber would require a spar capsectional weight equal to at least 65% of the total blade sectionalweight, which is doable. This table demonstrates that the rotorscontemplated in the OSTR and OSR patents must have used high moduluscarbon fibers, such as Toray™ M55J™. Because of the deficiencies oflaminates using high modulus carbon fibers, including fragility(resulting from a range of 34%-42% elongation to failure of mediummodulus carbon fibers), low compressive stress allowables with open holeand compressive stress allowables after impact of the laminate, and highcost, there is still a need to develop performance equivalent rotorsusing medium modulus fibers.

In this current CIP application, rotors have been designed using widerblades, of lower taper ratio, to provide the stiffness and massproperties required for high performance OSTR rotor blades.

FIG. 22 is a table showing details of the new rotor blade that usesmedium modulus fibers, and has a wider root chord and a lower taperratio than previously disclosed.

FIG. 23 is a cross-section of the rotor blade of FIG. 22 at station 40%of blade radius. The cross section in FIG. 23 depicts a multi-cellstructure with a leading edge cell, three box-shaped cells, and atrailing edge section. The four forward cells, including the leadingedge, are a co-cured structure containing a distributed cap at the outermold line of the airfoil shape; 80-90% of this cap is comprised ofcomposite plies with a fiber direction between −15 degrees and +15degrees relative to the blade span, with most of the fibers aligned at 0degrees. The remaining 10-20% of the cap is comprised of composite plieswith a fiber direction of +45 and −45 degrees relative to the bladespan. Inward of the outer cap is a cored section for panel stiffness,supported on the inner mold line by a carbon fiber composite face sheetwith fibers aligned +75 and −75 degrees relative to the blade span. Thetrailing edge, in contrast, has a thin outer laminate comprised offibers at +45 and −45 fiber direction, a core section, and a thin innerlaminate with fibers at +75 and −75 fiber direction.

FIG. 24 is a table showing sample carbon fiber flight thickness andorientation in cell C along the blade.

The carbon fiber composite cap has been distributed throughout the bladein the chordwise and spanwise directions to achieve the mass andstiffness properties shown in FIG. 25 mass and FIG. 26 stiffness withthe medium modulus fiber. This, in turn, achieves adequate deflectionresponse to aircraft in high lift loading when rotor-borne as shown inFIG. 27, the 5,000 lb lift per blade is equivalent to 20 psf discloading. This combination of blade planform shape, airfoil thickness,and carbon fiber cap distribution also achieves the needed aeroelasticstability with the medium modulus fibers. FIG. 28 presents a fan plot ofthe new blade, and demonstrates the high stiffness (first flap frequencyof 3.4 per rev, at the highest RPM of 350) of the new blade using themedium modulus fibers.

FIG. 29 shows that the new rotor design is free of aeroelasticinstability in hover, and FIG. 30 shows that the coupledrotor-nacelle-wing system is free of aeroelastic whirl instability inhigh speed forward flight beyond the aircraft dive speed of 450 Knots(120% of maximum cruise speed).

FIGS. 31 a-d is a collection of line drawings of various views of apreferred OSTR-equipped, twin tilt rotor VTOL aircraft offering highcruise efficiency, all drawn to scale, with a lift to drag ratio of 36.0at Mach number of 0.55, with a fuselage cross-sectional area that is5.25% of the total wing area. The aircraft also offers 32.0 lift to dragratio at Mach number 0.6. FIG. 32 provides the tabulated data for thataircraft.

FIG. 33 presents variations of lift to drag ratio of the same aircraftat Mach number of 0.55, with different fuselage sizes, varying thecross-sectional area in the range of 3%-15% of the wing area, at twodifferent variations of the fuselage lengths. These variations establishdifferent lift to drag (glide) ratios from approximately 26 to over 40.

Other Aspects

One skilled in the art will appreciate that the teachings herein can beviewed from many other aspects.

Viewed from one such additional aspect, preferred embodiments ofcontemplated aircraft comprise a high efficiency tilting rotor and wingdesign that enable both vertical takeoff and efficient cruising suchthat the aircraft can be commercially competitive with runway dependentaircraft operating in a range of 100 to 1000 or more miles.

Commercially viable embodiments are likely to have at least two tiltingrotors, and a wing design that both supports the rotors and is highlyefficient. In especially preferred embodiments, the wings havesufficient aspect ratio and wing area to maximal fuselage crosssectional area ratio to provide a cruise lift-to-drag ratio of 13 to 26or even 15 to 30.

Also, improvements to efficiency include one or more of a wing sweep ofless than 15°, 13°, or even less than 10°, and a thickness ratio of lessthan 26%, 23% or 21%, configured in a combination sufficient to achievea cruise lift-to-drag ratio of least 13 at a Mach number of 0.6.Combinations are especially contemplated that achieve that desiredlift-to-drag ratio at Mach number 0.65.

Viewed from another aspect, an inventive propulsion system of anaircraft is designed to be efficient in both a hover mode of theaircraft and in a cruise mode of the aircraft. This propulsiveefficiency is advantageously combined with a vertical takeoff emptyweight fraction could be less than 0.55, less than 0.50, or even lessthan 0.45. The propulsion system can advantageously comprise a tiltingrotor structured to have a hover Figure of Merit of at between 0.83 and0.90 at a disc loading of at most 20, 30, 40, 50, 60 or 70pounds-per-square-foot at sea level and in standard atmosphericconditions. Furthermore, the tilting rotor can be advantageouslystructured to a have cruise propulsion efficiency at between 0.82 and0.92 at a flight speed of Mach 0.6 and cruise altitude while producingat most 5% of the maximum rotor hover thrust in hover at sea level andin standard atmospheric conditions.

Prior art tilt-rotors can operate in the range from 83 to 100% ofmaximum rpm. Preferred propulsion systems according to the presentsubject matter can comprise an engine and drive train that operates therotor in cruise flight at an rpm of at most 70% of the maximum rpm inhover, in some cases 40-45%, 45-50%, 50-55%, 55-60% 60-65% and 65-70%.

The tilting rotor blade root thickness is considered to have asignificant effect on cruise efficiency at high Mach. In preferredembodiments, the rotor blade root thickness ratio is less than 25%, 22%or even 20%. This can yield a cruise propulsion efficiency of at least0.85, at least 0.88 and at least 0.90. A corresponding hover Figure ofMerit can realistically be at least 0.85, at least 0.87 and at least0.89.

Viewed from yet another aspect, an aircraft having a cruise effectivelift-to-drag ratio of at least 13 at a flight Mach of 0.50, or even 0.60(which historically has precluded vertical take off aircraft), can beimproved by adding rotors that lift the aircraft.

Viewed from still another aspect, methods of providing a verticaltakeoff transport aircraft with improved normalized cruise fuelproductivity comprise providing the aircraft with an efficient tiltingrotor, a variable speed engine and drive train, and a low empty weightfraction. In preferred embodiments, the normalized cruise fuelproductivity can be between 5.0 and 12.0, and more preferably 7.0 and12.0 (tons of payload x nautical miles of range divided by pounds offuel burned divided by engine specific fuel consumption expressed inpounds of fuel burned per horsepower per hour).

Additional methods for providing a vertical takeoff transport aircraftwith improved normalized cruise fuel productivity include providing aninboard wing supporting the tilting rotors. The wing can advantageouslybe mechanized by constructing it with sweep of less than 5°, 10°, oreven 15° in combination with a thickness ratio of less than 20%, 21%,23%, or even 25%. Such a wing can be constructed to have a planform ofsize and shape in a combination sufficient to achieve a cruiselift-to-drag of least 11, 12, 13, 14 or 15 at a Mach number of 0.6 oreven 0.65. One can also advantageously use an outboard wing to improvenormalized cruise fuel productivity by increasing total wing liftingarea, and thus the aircraft total lift/drag ratio. One could structuresuch an aircraft efficiently enough such that it could realisticallyhave a maximal fuselage cross sectional area larger than 8%, 10%, oreven 12% of wing area.

Normalized cruise fuel productivity can also be increased by efficientuse of composite materials, and independently by: (a) employing a lowerweight gearbox; (b) using a spinnion to replace a wing spar and atrunnion, and in other ways eliminating redundant structure, and (c)making other structural changes that reduce the empty weight of anaircraft. Preferred embodiments combine such methods to reduce the emptyweight fraction of the aircraft to 55, 50, 45, or even 40%.

Viewed from another aspect, an aircraft can be designed to be highlyefficient by an appropriate selection of aircraft components and byutilizing advanced design and analysis techniques, which allow theaccurate prediction of an aircraft's physical behavior.

In addition, the aircraft can be designed to support a tilting rotor bya wing that is sized and dimensioned to support the aircraft at themaximum hover weight in cruise-mode flight at an altitude of 25,000,30,000, 35,000, 40,000, or even 45,000 feet. This wing could be designedto have a geometric shape to provide a high maximum lift limit, orM²C_(L) limit, at a high cruise speed. It is contemplated that preferredembodiments could achieve a M²C_(L) limit of 0.30, 0.35, or even 0.40.It is envisioned that the wing can further be designed to have thintransonic wing airfoil shapes, of at most 25%, 24%, 23%, 22%, 21%, oreven 20%, which would assist in achieving a wing section lift to dragratio greater than 35, 45, or even 55 at a Mach number of 0.60, 0.61,0.63, 0.65, or even 0.67.

Still further, the aircraft can be designed to drive the tilting rotorusing a variable speed drive system. The drive system is preferablydesigned to be capable of operating the rotor at a cruise rpm of at most40-45%, 45-50%, 50-55%, 55-60% 60-65% and 65-70%, 70-75% and 75-80% ofthe maximum rpm in hover.

The design of such an aircraft can be advantageously facilitated usingcomputer simulation and numerical optimization, to modify the bladegeometry of a tilt rotor to provide improved hover figure of merit andcruise propulsive efficiency. It is further contemplated that thecomputer simulation could numerically solve unsteady Navier-Stokes fluidflow equations on a finite volume grid around geometry representative ofthe aircraft and rotors.

Independently, it is envisioned that the computer simulation andgeometry modification procedure could model any or all of the bladestructural dynamics, blade deflection, and/or blade aerodynamics,including effects of the rotor trailed wake, wing structural dynamics,wing deflection, wing aerodynamics, and the interactional effects ofthese elements. One of ordinary skill in the art could utilize theteachings herein to advantageously configure the numerical optimizationprocedure to comprise reduction of a plurality of objective functions,wherein the objective functions comprise performance metrics indifferent flight modes and conditions.

It is also contemplated to use computer simulation and numericaloptimization methods to modify a tiltrotor nacelle spinner geometry in amanner that improves aircraft efficiency by locally lowering Machnumbers at the inboard portion of the rotor blades. It is still furtherenvisioned that such computer simulation could be configured tosimultaneously compute aerodynamic effects of any or all of the wing, amechanically coupled nacelle, and a mechanically coupled spinner on ablade of a tilt rotor operated by a propulsion system.

Viewed from a different aspect, a rotorcraft having multiple rotorscapable of hovering the rotorcraft with disc loading between 10 and 70pounds per square foot could be improved by adding a wing to support therotors that also provides lift in forward flight. In preferredembodiments, the wing size and geometry is chosen to allow the aircraftto realistically cruise at a speed of Mach 0.6, 0.65, or even 0.67 at analtitude of at least 25,000, 30,000, 35,000, 40,000, or even 45,000 feetwith a cruise lift-to-drag ratio of at least 11, 13, 15, 17, 19, 21, 23,25, or even 27.

It should be apparent to those skilled in the art that many moremodifications besides those already described are possible withoutdeparting from the inventive concepts herein. The inventive subjectmatter, therefore, is not to be restricted except in the spirit of theappended claims. Moreover, in interpreting both the specification andthe claims, all terms should be interpreted in the broadest possiblemanner consistent with the context. In particular, the terms “comprises”and “comprising” should be interpreted as referring to elements,components, or steps in a non-exclusive manner, indicating that thereferenced elements, components, or steps may be present, or utilized,or combined with other elements, components, or steps that are notexpressly referenced. Where the specification claims refers to at leastone of something selected from the group consisting of A, B, C . . . andN, the text should be interpreted as requiring only one element from thegroup, not A plus N, or B plus N, etc.

What is claimed is:
 1. A tilt rotor aircraft, capable of hover takeoff, comprising: a fixed wing; a fuselage having a maximum cross sectional area 3-15% of a planform area of the fixed wing; a rotor coupled to the fuselage through the wing, and having at least first and second radially extending rotor blades, an engine for providing power to rotate the rotor; the fuselage, wing, and rotor configured to provide a glide ratio of 26-40 at 0.55 Mach.
 2. The aircraft of claim 1, wherein the wing and rotor are configured to have a ratio of wing span to rotor diameter of between 2.5 and 3.5.
 3. The aircraft of claim 2, wherein the rotor has a diameter of at least 15 feet.
 4. The aircraft of claim 2, wherein the rotor has a diameter of at least 25 feet.
 5. The aircraft of claim 1, wherein the glide ratio at Mach number 0.6 is not less than 85% of the glide ratio at Mach number 0.55.
 6. The aircraft of claim 1, wherein a maximum cross sectional area is 3-8 of the planform area of the fixed wing, and the fuselage, wing, and rotor are configured to provide the glide ratio of 32-40 at 0.55 Mach.
 7. The aircraft of claim 1, wherein the aircraft has a rotor maximum RPM in helicopter mode, and the fuselage, wing, rotor and engine are configured to provide flight operationability in airplane mode at variable RPM speeds below 70% of the rotor maximum RPM.
 8. The aircraft of claim 1, wherein the rotor has a disc loading no higher than 60 psf. 